An aerofoil portion of the turbine blade and the platform are heated to high temperature by high-temperature combustion gas flowing in a gas turbine. This causes the aerofoil portion and the platform to thermally expand outward in a radial direction of a rotor. As the aerofoil portion and the platform thermally expand at different rates, the heat elongation of the aerofoil portion and the platform causes heat stress between a hub of the aerofoil portion and the platform connected to the hub. The heat stress acts intensively on a trailing-edge end of the hub, which tends to cause a crack in the trailing-edge end. Therefore, it is necessary to reduce the heat stress while suppressing the temperature increase in the aerofoil portion and the platform.
JP2001-271603A proposes, as shown in FIG. 10, to provide cooling channels 61 through 64 in the aerofoil portion 12 and the platform 60 and to form a concave 20 in a trailing-edge end part 22 of the platform 60 along a circumferential direction of the rotor (in a direction of passing through a plane of paper of FIG. 10). In the aerofoil portion 12, the cooling channels 61 to 63 are formed along the radial direction of the rotor from a base portion 2 through the aerofoil portion 12. In the platform 60, the cooling channel 64 is formed along the axial direction of the rotor from the trailing-edge end surface 18 to a leading-edge end portion of the platform 60. By streaming cooling air in the aerofoil portion 12 and the platform 60, the temperature increase of the aerofoil portion 12 and the platform 60 is prevented.
Further, in response to the heat elongation of the aerofoil portion 12 expanding outwardly in the radial direction of the rotor, the trailing-edge end surface 18 disposed outside of the concave 20 in the radial direction of the rotor, expands outwardly in the radial direction of the rotor. By this, concentration of the heat stress on the trailing edge end portion 22 of the hub 13 is prevented.